The present invention relates to an air cooling system for supplying cooling air to the turbine of a gas turbine engine.
The performance of an aircraft gas turbine engine is directly related to the maximum permissible temperature of the combustion gases. The amount of heat that the high pressure turbine components may withstand is generally the limiting factor on the upper limits of the gas temperature. It is well-known to provide cooling air to the turbine components located in the flow path of the high temperature gases, such as the rotor wheel blades and the stationary vanes, in order to raise the temperature of the high temperature gases, without exceeding the structural limitations of the turbine components.
The conventional practice in the known high pressure turbine cooling systems is to tap a portion of the air from the high pressure compressor and direct it onto the turbine blades and/or vanes. To accelerate the cooling air, accelerator means are provided in the air tapping zone. The air from the high pressure compressor is tapped in a centripetal manner through a rotor disk of the high pressure compressor with the accelerator means being located inside the rotor disk. The air is directed through a tube, conduit or passageway extending generally along the drive shaft interconnecting the turbine rotor with the compressor rotor. The air is then directed in a centrifugal manner onto the stationary vanes and the moving blades of the gas turbine.
Typical accelerating means for such cooling systems can be found in French Patent Applications A2,552,164; 2,609,500 and 2,614,654. In these systems, the accelerating means are located in the rotor disk of the high pressure compressor.
In British Patent 2,189,845, a system is disclosed for moving cooling air to a turbine which includes a centrifugal nozzle associated with a rotor wheel stage of the turbine.
While such systems have proven beneficial when used with today's turbojet engines, they are inadequate to cool the next generation gas turbine engines, especially air-breathing, turbojetramjet engines. The extremely high air temperature in the high pressure compressors of these gas turbine engines effectively precludes the direct use of this air as a cooling medium. Due to its initially high temperature, the cooling air must be cooled in a heat exchanger before directing it onto the turbine structure.
French Patent Application A2,400,618 discloses a method for air cooling the high pressure turbine of a gas turbine engine in which the cooling air is tapped from the combustion chamber and cooled in a heat exchanger located outside of the gas turbine engine envelope in an air-bypass duct. The cooled air moves into rear risers of the high pressure compressors rear support and then into an expansion valve, before being directed onto the gas turbine structure. The circuitous path of the cooling air with resultant pressure losses will prevent the rapid air circulation required by the future high pressure gas turbine engines.